DE102004032978A1 - Flow structure for a turbocompressor - Google Patents

Abstract

The invention relates to a flow structure for a turbocompressor. DOLLAR A The flow structure comprises an annular chamber (18) which is arranged concentrically to an axis (17) of the turbocompressor in the region of free blade ends of a blade ring (15) and which radially adjoins a main flow channel (12), wherein the annular chamber (18) is bounded by a front, upstream wall (19), a rear, downstream wall (20) and a substantially axially extending wall (21), wherein in the annular chamber (18) guide elements (22) are arranged and wherein the annular chamber (18 ) in a front and / or rear area allows a flow passage in the circumferential direction. DOLLAR A According to the invention in the region of the substantially axially extending wall (21) or in the region of the upstream wall (19) an opening (30) is arranged, which allows a flow passage out of the annular chamber (18) out, wherein at least one compression chamber (31). is present for receiving this emerging flow.

Description

The The invention relates to a flow structure for a Turbo compressor according to the preamble of claim 1. Des Furthermore, the invention relates to a turbocompressor and a Aircraft engine and a stationary one Gas turbine.

flow structures or circulation structures for Turbo compressors are known in the form of so-called "Casing Treatments" and "Hub Treatments". The "Casing Treatments" and "Hub Treatments" flow structures have primarily the task the aerodynamically stable operating range of the compressor to increase an optimization of the surge margin. An optimized surge margin allows higher Compressor pressures and thus a higher one Compressor load. The for a local stall and ultimately for that Pumping the compressor responsible faults occur on the housing side Ends of the blades of one or more compressor stages or on the hub-side, radially inner ends of the vanes because, in these areas, the aerodynamic load in the compressor is highest. Through flow structures or circulation structures, the flow in the area of the blade ends stabilized. On the one hand, such circulation structures permit one another a flow in the axial direction and on the other hand, a flow passage in the circumferential direction. flow structures in the area of the housing side Ends of the blades are referred to as "casing treatments", flow structures in the area the hub-side ends of the vanes are referred to as "hub treatments".

The DE 33 22 295 C3 discloses an axial fan with a circulation structure. The axial fan according to DE 33 22 295 C3 comprises an annular chamber, which is arranged concentrically to an axis of the axial fan in the region of free ends of a blade ring, wherein the annular chamber is radially connected to a main flow channel. In the annular chamber guide elements are fixed. The circulation structure according to this prior art allows a flow in the axial direction and in the circumferential direction. Characteristic of this type of zirkulationsstrukturen is a aligned with the contour of the main flow channel, closed ring, which separates the rear entrance region of the circulation structure from the front exit region thereof and forms a smooth, closed surface region.

The US 5,282,718 discloses a "Casing Treatment" with arranged in an annular chamber vanes. In this "casing treatment" as well, the inlet area and the outlet area of the circulation structure are separated by a solid ring which is smooth and closed towards the blades. Such rings in the blade area must be provided in the event of contact with the blade tips with a Anstreif- or inlet lining.

Another "Casing Treatment" is from the US 5,137,419 known. In the disclosed therein circulation structure to improve the surge margin of a compressor in Axialbauweise axial or axially oblique grooves are present. In the circulation structure disclosed therein, however, no flow passage in the circumferential direction is possible.

Further circulation structures for turbocompressors are from the US 4,511,308 as well as the DE 35 39 604 C1 known.

Farther it is already state of the art, in the compressor area between adjacent vane rings and blade wreaths through openings in the compressor housing - through so-called "Bleed Slots "- to divert air, which is used for example as cabin air or as cooling air for the turbine. As a result, however, the flow conditions at the downstream lying blade lattice in the sense of reducing the surge margin negatively influenced.

Of these, Based on the present invention, the problem underlying a novel flow structure for one To create turbocompressors, in particular a branching of Compressor air without fluidic Disadvantage allows.

This Problem is solved by that the initially mentioned flow structure by the features of the characterizing part of the claim 1 is further developed. According to the invention, at least in the area one of the walls at least one opening arranged, which has a flow passage out of the annular chamber allows, where at least one compressor chamber for receiving the exiting air is available.

The inventive flow structure acts at least partially as a casing treatment or stroke treatment with all of them Advantages and allows a fluidic optimized branching of air. The invention thus enables on the one hand air circulation and on the other hand a branch of Air. If no air is diverted, it acts exclusively as Casing Treatment or Hub Treatment.

According to an advantageous embodiment of the invention, the guide elements limit on the one hand as in the axial direction extending flow channels and on the other hand, at least one circumferentially extending flow channel, wherein the or each opening is disposed at an upstream end of the axially extending flow channels.

Preferably is a medium, namely Air, which over the flow channels running in the axial direction and over the or each opening in the Compressor chamber flows in over one outlet opening discharged from the compression chamber.

Of the Turbo compressor according to the invention is defined by the features of claim 17, the aircraft engine according to the invention is in claim 18 and the stationary gas turbine according to the invention is in claim 19 defined.

preferred Further developments of the invention will become apparent from the dependent claims and the following description.

embodiments The invention will be described, but not limited to, with reference to the drawing explained in more detail. In the drawing shows:

1 a partial longitudinal section through an axial compressor in the region of a housing-side flow structure according to the invention;

2 a partial cross section through the arrangement according to 1 along cutting direction II in 1 ;

3 a partial cross section through the arrangement according to 1 along cutting direction II-II in 1 ;

4 : A partial longitudinal section through a compressor in axial construction in the region of a hub-side flow structure according to the invention.

Hereinafter, the present invention will be described with reference to FIG 1 to 4 described in more detail.

1 shows a partial longitudinal section through a compressor 10 in axial construction, with a housing 11 of the compressor 10 a main flow channel 12 limited. The flow direction through the main flow channel 12 is in 1 through an arrow 13 visualized.

In the main flow channel 12 are arranged in the axial direction one behind the other several fixed vanes and a plurality of rotating blades. This is how the section from the compressor shows 10 according to 1 a first vane ring 14 , wherein in the flow direction behind this first vane ring 14 a blade ring 15 and in the flow direction behind the blade ring 15 again a vane ring 16 in the main flow channel 12 is arranged. The rotating blades of the rotor blade ring 15 rotate together with a rotor opposite the fixed housing 11 as well as against the fixed vanes of the vane rings 14 and 16 , An axis of rotation of the rotor with which the blades rotate is in FIG 1 with the reference number 17 characterized.

To the main flow channel 12 borders according to 1 in the radial direction an annular chamber 18 at. The annular chamber 18 is from an axially front wall 19 , an axially rear wall 20 and a radially outer wall 21 limited. The axially front wall 19 and the axially rear wall 20 extend substantially in the radial direction. The radially outer wall 21 runs essentially in the axial direction. The annular chamber 18 is in the embodiment shown, the housing-side ends 35 the blades of the rotor blade ring 15 assigned. It should be noted that such an annular chamber also hub-side, radially inner ends of the vanes of the vane rings 14 and 16 can be assigned.

In the ring chamber 18 are guiding elements 22 arranged. The representation according to 2 it can be seen that the guiding elements 22 at an angle β to the blades of the blade ring 15 or are inclined relative to a radial. The representation according to 2 shows the guide elements in the axial direction. In the circumferential direction are the same by the angle β with respect to the blades of the blade ring 15 inclined. The angle β is chosen so that a flow outlet from the main flow channel 12 in the ring chamber 18 is facilitated in terms of flow and therefore takes place without major losses. The direction of rotation of the blade ring 15 is in 2 with the arrow 23 characterized. It should be noted at this point that the inclination angle β does not have to be constant in the circumferential direction, as in the exemplary embodiment shown, but rather can change in the circumferential direction. The guide elements then have a curved course in the circumferential direction. The angle of inclination β can decrease from radially inward to outward to a value "zero", with appropriately curved guide elements. In any case, the angle β is chosen so that the flow enters the annular chamber 18 Optimized. As 2 can be taken, are the guiding elements 22 not only characterized by the angle β but also by their radial height H and their distance in the circumferential direction, which is defined by the parameters T and P.

The guiding elements 22 are scoop-like and have a spatially curved profile with varying thickness. This can in particular the representation according to 3 be removed. 3 shows the profiling of the guide elements 22 and the blades of the blade ring 15 , where the direction of rotation of the blade ring 15 again by the arrow 23 is visualized. In 3 It can be seen that the flow outlet in the region of an upstream edge 24 the annular chamber 18 in relation to the blade ring 15 to be done with counter-rotation. A downstream edge of the annular chamber 18 is in 3 with the reference number 25 designated.

As 1 can be removed, the guide elements extend 22 within the annular chamber 18 not over the entire axial length of the annular chamber 18 rather, leave the vanes 22 in the area of the axially rear wall 20 - So at the downstream edge 25 - a recess 26 free. The guiding elements 22 thus limit a plurality of flow channels extending in the axial direction 27 and a circumferentially extending flow channel 28 , Within the annular chamber 18 Accordingly, in addition to the predominantly extending in the axial direction recirculation flow in the circumferential direction is possible. The axial recirculation is in 1 through the arrow 29 characterized.

At this point it should be noted that in contrast to the embodiment shown in the annular chamber and guide elements can be arranged, both in the region of the axially front wall 19 as well as in the area of the axially rear wall 20 Release recesses that allow flow passages in the circumferential direction. In the embodiment of 1 is only in the area of the axially rear wall 20 such a recess 26 intended. Furthermore, it is within the meaning of the invention possible that the defined by the guide elements recess to ensure a flow passage in the circumferential direction exclusively in the region of the axially front wall 19 is positioned.

1 can be seen that the main flow channel 12 facing ends of the guide elements 22 on a contour of the main flow channel 12 run and axial with free ends 35 the blades of the rotor blade ring 15 overlap. The main flow channel 12 facing ends of the guide elements 22 have such a distance from the free ends in the radial direction 35 the blades of the rotor blade ring 15 on that in normal operation of the turbocompressor the guide elements 22 the free ends 35 the blades of the rotor blade ring 15 do not touch. It should be noted at this point that the guiding elements 22 may consist of a metal, such as steel, or a nickel-based alloy or a cobalt-based alloy. Furthermore, it is possible, the guide elements 22 from a light metal, such as aluminum, or a plastic, such as thermoplastics, thermosetting plastics or elastomers. The annular chamber 18 can be including the guide elements 22 in one piece or as at least two circumferentially adjacent segments are executed. In the case where the annular chamber 18 as well as the guiding elements 22 when two segments are executed, are the guiding elements 22 detachable in the annular chamber 18 attached.

In the sense of the present invention is in a main flow direction 16 seen upstream region of the annular chamber 18 at least one opening 30 arranged, which has a flow passage from the annular chamber 18 out into a compactor chamber 31 allows. In the embodiment of 1 is the opening or recess 30 in the substantially axially extending wall 21 the annular chamber 18 integrated. The opening 30 is in an area of the radially outer wall 21 introduced, which is attached to the axial front wall 19 followed. The or each opening 30 in the region of the wall extending substantially in the axial direction 21 is therefore arranged at an upstream end. The or each opening 30 in the region of the wall extending substantially in the axial direction 21 the annular chamber 18 allows flow passage in the radial direction from the annular chamber 18 out into a compactor chamber 31 , The compressor chamber 31 can also be referred to as compressor pre-chamber. The flow passage from the annular chamber 18 in the compression chamber 31 is in 1 through the arrow 32 characterized. The flow in the area of the annular chamber 18 on the one hand serves the recirculation back into the main flow channel 12 and, on the other hand, a circulation from the main flow channel 12 out into the compressor chamber 31 ,

In contrast to the embodiment shown, it is also possible the or each opening 30 in the axial front wall 19 contribute. Even then, the opening is in an upstream region of the annular chamber 18 arranged.

According to 1 is that from the ring chamber 18 in the compression chamber 31 flowing medium via an outlet opening 33 from the compression chamber 31 out laxable, what in 1 through an arrow 34 is designated. The compressor chamber 31 closes radially to the annular chamber 18 and surrounds the annular chamber 18 radially outside.

In connection with the embodiment of the 1 to 3 It should be noted at this point that the axial center of the annular chamber 18 upstream of the axial center of the associated blade ring 15 lies. By this is meant that the axial center of the annular chamber 18 on the associated wall of the main flow channel 12 upstream of the axial center of the free ends 35 the blades of the rotor blade ring 15 lies.

With reference to 1 to 3 the flow structure according to the invention has been described using the example of a ring chamber forming a casing treatment. As already mentioned, the flow structure according to the invention can also be used in the region of the hub-side ends of stationary vanes of a vane ring to form a so-called stroke treatment. So shows 4 a partial longitudinal section through a compressor, in which the flow structure according to the invention in the region of a hub of a compressor rotor is designed as a hub treatments.

4 again shows a compressor 10 in Axialbauweise, wherein in the main flow channel 12 of the compressor 10 in the axial direction one behind the other several fixed vane rings and a plurality of rotating blade rings are arranged. 4 shows in this regard a blade ring 15 , wherein in the flow direction behind this blade ring 15 a vane ring 16 and in the flow direction behind the vane ring 16 turn a blade ring 15 is positioned. The main flow channel 12 is hub side of an outer contour of a hub 36 limited.

In contrast to the embodiment of 1 to 3 is in the embodiment of 4 the ring chamber 18 not radially outside the housing side but much more radially formed on the hub side. The annular chamber 18 is free ends 37 the fixed vanes of the vane ring 16 assigned. As for the other details, the hub-side flow structure of the embodiment of 4 with the housing-side flow structure of the embodiment of 1 to 3 the same reference numerals are used to avoid unnecessary repetitions for the same assemblies. It may refer to the comments on the embodiment of 1 to 3 to get expelled.

aid the flow structure according to the invention for one Turbo compressor is therefore on the one hand a recirculation within of the main flow channel and on the other hand, a circulation from the main flow channel out into a compressor chamber possible. This is done by the formed according to the invention Ring chamber allows. The flow structure according to the invention causes an optimization of the surge margin of the turbocompressor. It is both in turbocompressors in Axialbauweise and in Turbo compressors can be used in diagonal construction or radial construction. The flow structure according to the invention is preferably used in aircraft engines or stationary gas turbines.

10

compressor

11

casing

12

Main flow channel

13

arrow

14

vane ring

15

Blade ring

16

vane ring

17

axis

18

annular chamber

19

wall

20

wall

21

wall

22

vane

23

arrow

24

edge

25

edge

26

recess

27

flow channel

28

flow channel

29

arrow

30

opening

31

compression chamber

32

arrow

33

outlet opening

34

arrow

35

The End

36

hub

37

The End

Claims (19)

Flow structure for a turbocompressor, in particular for a compressor of a gas turbine, with at least one annular chamber ( 18 ) concentric with an axis ( 17 ) of the turbocompressor in the area of free blade ends ( 36 ; 37 ) of a blade ring ( 15 ; 16 ) and arranged radially to a main flow channel ( 12 ), wherein the annular chamber ( 18 ) from a front, upstream wall ( 19 ), a rear, downstream wall ( 20 ) and a substantially axially extending wall ( 21 ) is limited, wherein in the annular chamber ( 18 ) Guiding elements ( 22 ) are arranged, and wherein the annular chamber ( 18 ) in a front and / or rear area allows a flow passage in the circumferential direction, characterized in that in the region of at least one of the walls ( 19 . 20 . 21 ) at least one opening ( 30 ) is arranged, which has a flow passage from the Annular chamber ( 18 ) and that at least one compression chamber ( 31 ) is present for receiving this exiting flow. Flow structure according to claim 1, characterized in that in the substantially axially extending wall ( 21 ) of the annular chamber ( 18 ) in an upstream region of the same at least one opening ( 30 ), wherein the or each aperture ( 30 ) a flow passage in the radial direction from the annular chamber ( 18 ) out into the compression chamber ( 31 ). Flow structure according to claim 1 or 2, characterized in that the compression chamber ( 31 ) an outlet opening ( 33 ), via which into the compression chamber ( 31 ) from the annular chamber ( 18 ) incoming air from the compression chamber ( 31 ) is dissipatable. Flow structure according to one or more of claims 1 to 3, characterized in that the compression chamber ( 31 ) radially to the annular chamber ( 18 ) adjoins. Flow structure according to one or more of claims 1 to 4, characterized in that the annular chamber ( 18 ) in the area of the housing-side ends ( 35 ) of moving blades ( 15 ) is arranged, and that the compression chamber ( 31 ) the annular chamber ( 18 ) encloses radially on the outside. Flow structure according to one or more of claims 1 to 5, characterized in that the annular chamber ( 18 ) in the area of hub-side ends ( 37 ) of vanes ( 16 ) is arranged, and that the compression chamber ( 31 ) the annular chamber ( 18 ) encloses radially inside. Flow structure according to one or more of claims 1 to 6, characterized in that the guide elements ( 22 ) on the one hand in predominantly axial direction extending flow channels ( 27 ) and on the other hand, at least one circumferentially extending flow channel ( 28 ), whereby the or each opening ( 30 ) at an upstream end of the axially extending flow channels ( 27 ) in the substantially axially extending wall ( 21 ) is arranged. Flow structure according to one or more of claims 1 to 7, characterized in that the guide elements ( 22 ) in the axial direction in the circumferential direction by an angle (β) are inclined or curved in the circumferential direction. Flow structure according to claim 8, characterized in that the angle (β) is chosen so that the flow inlet from the main flow channel ( 12 ) in the annular chamber ( 18 ) is facilitated fluidically. Flow structure according to one or more of claims 1 to 9, characterized in that the main flow channel ( 12 ) facing ends of the guide elements ( 22 ) on and / or near a contour of the main flow channel ( 12 ) and axially overlap with the free blade ends or axially adjacent to the region of the free blade ends. Flow structure according to claim 10, characterized in that the main flow channel ( 12 ) facing ends of the guide elements ( 22 ) have in the radial direction such a distance from the free blade ends of the turbocompressor that during normal operation of the turbocompressor the guide elements ( 22 ) do not touch the free blade ends. Flow structure according to one or more of claims 1 to 11, characterized in that the guide elements ( 22 ) are scoop-like, spatially curved, with varying thickness and defined profile sections are executed. Flow structure according to one or more of claims 1 or 12, characterized in that the annular chamber ( 18 ), the guiding elements ( 22 ) and the compression chamber ( 31 ) in a housing ( 11 ) of the turbocompressor are integrated. Flow structure according to one or more of claims 1 or 13, characterized in that the annular chamber ( 18 ), the guiding elements ( 22 ) and the compression chamber ( 31 ) in a hub ( 36 ) of the turbocompressor are integrated. Flow structure according to one or more of claims 1 or 14, characterized in that the axial center of the annular chamber ( 18 ) upstream of the axial center of the associated blade ring ( 15 ; 16 ) lies. Flow structure according to claim 15, characterized in that the axial center of the annular chamber ( 18 ) at the associated main flow channel wall upstream of the axial center of the free blade ends of the associated blade ring ( 15 ; 16 ) lies. Turbo compressors in axial design and / or diagonal design and / or radial construction, with at least one flow structure according to one or more of claims 1 to 16. Aircraft, with a turbocompressor according to claim 17th Stationary gas turbine, with a turbover Denser according to claim 17.